The present invention relates generally to a gas turbine engine cooling component for end rail cooling, and in particular a turbine engine shroud where each shroud segment provides cooling to both the high pressure and low pressure turbine sections of a gas turbine engine. The present invention further relates to a turbine engine subassembly, and in particular a shroud subassembly that uses a pair of such cooling segments in combination with at least one discourager and primary spline seal.
To increase the efficiency of gas turbine engines, a known approach is to raise the turbine operating temperature. As operating temperatures are increased, the thermal limits of certain engine components can be exceeded, resulting in material failure or, at the very least, reduced service life. In addition, the increased thermal expansion and contraction of these components adversely affects clearances and their interfitting relationships with other components of different thermal coefficients of expansion. Consequently, these components should be cooled to avoid potentially damaging consequences at elevated operating temperatures.
It is common practice then to extract from the main airstream a portion of the compressed air from the compressor for cooling purposes. So as not to unduly compromise the gain in engine operating efficiency achieved through higher operating temperatures, the amount of extracted cooling air should be held to a small percentage of the total main airstream. This requires that the cooling air be utilized with the utmost efficiency in maintaining the temperatures of these components within safe limits.
A particularly important component subjected to extremely high temperatures is the shroud located immediately downstream of the high pressure turbine nozzle, immediately downstream from the combustor. The shroud closely surrounds the rotor of the high pressure turbine and thus defines the outer boundary (flow path) of the extremely high temperature main (hot) gas stream flowing through the high pressure turbine. To prevent material failure and to maintain proper clearance with the rotor blades of the high pressure turbine, adequate shroud cooling is an important concern.
Shroud cooling can be achieved by impingement cooling of the back surface of the shroud, as well as cooling holes that extend from the back surface of the base of the shroud and through to the forward or leading edge of the shroud, the bottom or inner surface of the base in contact with the main (hot) gas stream, and the aft or trailing edge of the shroud to provide both convection cooling inside the holes, as well as impingement and film cooling of the shroud. Cooling flow is also provided through the side panels or rails as convection cooling inside the cooling passages or holes, as well as impingement cooling as cooling air exits from the holes. See, for example, commonly assigned U.S. Pat. No. 5,169,287 (Proctor et al), issued Dec. 8, 1992, which shows a prior embodiment of shroud cooling of the high pressure turbine section of a gas turbine engine. This cooling minimizes local oxidation and burning of the shrouds near the hot main or core gas stream in the high pressure turbine section. Indeed, the cooling holes that exit through the side panel of the shroud of commonly assigned U.S. Pat. No. 5,169,287 can provide important impingement cooling to the side panel of the adjacent shroud.
The leading edge of the shroud is subject to the hottest flow path gas or air, and has the highest heat transfer coefficient, making this section one of the most difficult to cool. As also shown in commonly assigned U.S. Pat. No. 5,169,287, a circumferential row of holes can be angled to also exit at the leading edge of the shroud to provide both convection and film cooling at the leading edge of the shroud. As this cooling film decays and mixes with the hot flow path air, additional circumferential rows of cooling holes can be required to provide more convection and film cooling.
Another type of shroud assembly for a different type of gas turbine engine is shown in commonly assigned U.S. Pat. No. 5,127,793 (Walker et al), issued Jul. 7, 1992. As shown particularly in FIGS. 4 and 4c of U.S. Pat. No. 5,127,793, this prior shroud assembly uses single-piece shroud segments 30 that are designed to span over both the high pressure and low pressure turbine sections of the gas turbine engine. As shown particularly in FIG. 4, cooling is provided by directing a portion of the cooling air 74 through ports 78 and through segmented impingement baffles 80 and against the high pressure portion 83 of shroud segment 30. Another portion of this air 74 is directed into cavity B, with most of it being delivered to cavity C located adjacent the low pressure portions 85 of each shroud segment 30 through holes 84 formed in the support cone portion 86 of turbine shroud support 44. An impingement baffle 81 attached to shroud support 44 directs and meters impingement cooling air from cavity C onto the low pressure portion 85 of shroud segment 30. While this prior shroud design of U.S. Pat. No. 5,127,793 provides significant impingement cooling to the back surface of shroud segment 30 in both the high and low pressure sections, it provides no impingement cooling to the side panels or rails of adjacent shroud segments.
The shroud assembly shown in commonly assigned U.S. Pat. No. 5,127,793 extends from approximately the aft end of the upstream turbine nozzle to approximately the leading edge of the downstream turbine nozzle and encloses (i.e., provides a 360xc2x0 annular structure around) the outer air flow path of a gas turbine engine that typically has a turning nozzle to direct the air flow properly into the blade row, then into a row of blades in the HPT section, and then into another row of blades in the LPT section. Axial gaps between these shroud segments allow for thermal growth over the large range of temperatures the gas turbine engine produces. As hot flow path air passes through the row of turbine blades, work is extracted from the air, thus creating a pressure and temperature drop axially through the blade row. As a result, both the pressure and temperature is higher at the leading edge of the shroud and lower at the trailing edge of the shroud.
A typical sealing method along the axial split lines or gaps between shroud segments is to provide a machined groove or slot in which a thin metal seal (usually referred to as a xe2x80x9cspline sealxe2x80x9d) is placed, with pressure loading across the seal to provide positive sealing and to minimize air leakage. See FIG 11a of commonly assigned U.S. Pat. No. 5,127,793 which shows a pair of longitudinally extending slots in shroud segment 30, the lower slot receiving the lower or xe2x80x9cdiscouragerxe2x80x9d spline seal, the upper slot(s) receiving the upper or xe2x80x9cprimaryxe2x80x9d spline seal(s). The portion of the axial segment gap that is set up between the shroud segments below the xe2x80x9cdiscouragerxe2x80x9d seal (commonly referred to as the xe2x80x9ctrenchxe2x80x9d) also has hot flow path air traveling axially down it due to the pressure gradient produced by the turbine blade row. Typically no preferential cooling is added to this xe2x80x9ctrench.xe2x80x9d Instead, in the past, air that leaks around the xe2x80x9cdiscouragerxe2x80x9d seal and the conduction from adjacent metal has been deemed sufficient to cool the axial split lines, i.e., at the side rails or panels of the shroud segments. However, in more recent gas turbine engines that operate at higher temperatures, it has been discovered that oxidation and loss (melting) of the parent material along the axial split-lines of shroud segments can occur.
Accordingly, it would desirable, therefore, to provide a shroud and resulting shroud assembly, particularly for the combined high pressure and low pressure turbine sections, that creates effective impingement cooling for the side panels of adjacent shroud segments. It would also be desirable to provide such impingement cooling while efficiently utilizing the total available cooling air so as not to significantly decrease the efficiency of the gas turbine engine. It would further be desirable to provide effective cooling and purging in the xe2x80x9ctrenchxe2x80x9d between the shroud segments that are below the xe2x80x9cdiscouragerxe2x80x9d seal.
The present invention relates to a turbine engine cooling component such as a shroud segment for a combined high pressure and low pressure turbine section of a gas turbine engine that provides effective end rail cooling to the side rails or panels of adjacent turbine cooling components (e.g., at the axial split lines between adjacent shroud segments), as well as effective cooling in the gap or xe2x80x9ctrenchxe2x80x9d between adjacent turbine engine cooling components (e.g., adjacent shroud segments) that is below the discourager spline seal. This turbine cooling component comprises:
(a) a circumferential leading edge;
(b) a circumferential trailing edge spaced from the leading edge;
(c) an arcuate base connected to the trailing and leading edges and having a back surface and an arcuate inner surface that is in contact with the main (hot) gas stream of the gas turbine engine moving in the direction from the leading edge to the trailing edge of the turbine component;
(d) a pair of spaced opposed axial side panels connected to the leading and trailing edges;
(e) each of the side panels having a lower discourager spline seal slot extending longitudinally from the leading edge to the trailing edge of each side panel that is capable of receiving an edge of a discourager spline seal, each lower slot having at least a bottom wall and a top wall;
(f) each of the side panels having an upper primary spline seal slot spaced above the lower slot and extending longitudinally from the leading edge to the trailing edge of each side panel that is capable of receiving an edge of a primary spline seal, each upper slot having at least a bottom wall and a top wall;
(g) a plurality of cooling air passages extending through the base from the back surface thereof and having spaced outlets exiting from at least one of the side panels between the bottom wall of the top slot and the bottom wall of the lower slot;
(h) a plurality of spaced air flow pathways along the length of the lower slot and below the bottom wall of the upper slot that are capable of receiving air flowing over and above the discourager seal when positioned in the lower slot and passing that air flow around the edge and beneath the discourager seal.
The present invention further relates to a turbine engine cooling subassembly comprising a pair of such adjacent turbine engine components, and having:
(1) opposed adjacent side panels having a gap therebetween and wherein the spacing of the air flow pathways along the length of the lower slot for each of the adjacent side panels is staggered such that the outlet of each of the cooling air passages exiting each adjacent side panel are opposite one of the air flow pathways of the other adjacent side panel;
(2) at least one discourager spline seal positioned in the gap between the opposed adjacent side panels and including a pair of spaced edges having a length and thickness such that each of the edges is capable of being received by the lower slot of one of the adjacent side panels;
(3) the at least one discourager seal being positioned below the outlet of each of the cooling air passages exiting each adjacent side panel;
(4) at least one primary spline seal positioned in the gap and including a pair of spaced edges having a length and thickness such that each of the edges is capable of being received by the upper slot of one of the adjacent side panels.
The turbine engine cooling component (e.g., shroud) of the present invention is particularly useful in providing effective, efficient and more uniform cooling to the end rail (i.e., split line) region, especially for the metal of the turbine component below the discourager seal. The turbine engine cooling subassembly (e.g., shroud cooling subassembly) of the present invention that comprises a pair of such turbine components (e.g., shroud segments) that have staggered or offset air flow pathways (preferably spaced recesses in the bottom wall of the lower slot) and outlets for the cooling air passages exiting from the adjacent side panels, also provide impingement cooling coverage to each of the adjacent side panels. In particular, this turbine cooling subassembly causes cooling air to flow: (a) over the discourager seal and then under it (via the air flow pathways such as the recesses in the bottom wall of the lower slot) to impinge on the side panel (below the lower slot) of the turbine component (e.g., shroud) from which the cooling air came from; (b) downstream above the discourager seal (via the air flow pathways) and out, such as through recesses in the bottom of wall the lower slot of same side panel from which the cooling air came to impinge on the adjacent side panel (below its lower slot); and (c) to purge the hot gas or air in the xe2x80x9ctrenchxe2x80x9d below the discourager seal.
The turbine engine cooling component of the present invention can also have certain optional but preferred features. One preferred feature is to have no cooling air passages exiting from certain portions of the side panel where cooling air flow is not required or is unnecessary, and thus economizes the usage of the total cooling air flow. Yet another preferred feature is to provide a sub-impingement pocket at the rear or aft portion of certain sections of the turbine cooling component, especially a shroud cooling segment having a high pressure turbine (HPT) section. This sub-impingement pocket helps reduce the source pressure of the cooling air provided to the rear or aft portion of the HPT section (which is usually at its lowest sink pressure in the HPT section) to provide an adequate amount of cooling air to the cooling air passages exiting from the side panel at the rear or aft portion of the HPT section and to reduce the total air flow emitting from such passages, again economizing the usage of the total cooling air flow.